Airfoil Comparison
Select airfoils from the airfoil database or add your own airfoils and compare the airfoil shape and lift/drag polars.
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(naca4412-il) NACA 4412 | NACA 4412 airfoil Max thickness 12% at 30% chord Max camber 4% at 40% chord | Remove Airfoil details Airfoil plotter |
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Polars for (naca4412-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca4412-il | 50,000 | 9 | 33.4 at α=9.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca4412-il | 50,000 | 5 | 36.1 at α=8.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca4412-il | 100,000 | 9 | 56.1 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca4412-il | 100,000 | 5 | 57.4 at α=6.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca4412-il | 200,000 | 9 | 78.1 at α=7° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca4412-il | 200,000 | 5 | 76.6 at α=6° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca4412-il | 500,000 | 9 | 107.5 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca4412-il | 500,000 | 5 | 101.1 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca4412-il | 1,000,000 | 9 | 129.4 at α=5.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca4412-il | 1,000,000 | 5 | 119.2 at α=4.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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